1. Field of the Invention
The invention relates generally to the field of thermal barrier coatings, and more particularly, to new thermal barrier coating compositions to extend the operating temperature capabilities for the components in a combustion turbine engine.
2. Background Information
The demand for continued improvement in the efficiency of combustion turbine and combined cycle power plants has driven the designers of these systems to specify increasingly higher firing temperatures in the combustion portions of these systems. Although nickel and cobalt based xe2x80x9csuperalloyxe2x80x9d materials are now used for components in the hot gas flow path, such as combustor transition pieces and turbine rotating and stationary blades, even these superalloy materials are not capable of surviving long term operation at temperatures sometimes exceeding 1,200xc2x0 C.
Examples of cobalt or nickel based superalloys are, for example, Cr.Al.Co.Ta.Mo.W, which has been used for making SC turbine blades and vanes for gas turbines, as taught, for example, in U.S. Pat. No. 5,716,720 (Murphy).
These turbine components are generally protected by a basecoat of MCrAlY, where M is selected from the group of Fe, Co, Ni, and their mixtures, as taught, for example, by U.S. Pat. Nos. 4,916,022; 5,238,752; 5,562,998; and 5,683,825 (Solfest et al.; Duderstadt et al.; Strangman; and Bruce et al., respectively). These basecoats are usually covered by an aluminum oxide layer and a final thermal barrier coating. The standard thermal barrier coating (xe2x80x9cTBCxe2x80x9d) is made from yttria-stabilized zirconia, ceria-stabilized zirconia, scandia-stabilized zirconia or non-stabilized zirconia, as taught, for example, by the Bruce et al. patent U.S. Pat. No. 5,683,825.
Much of the development in this field of technology has been driven by the aircraft engine industry, where turbine engines are required to operate at high temperatures, and are also subjected to frequent temperature transients as the power level of the engine is varied. A combustion turbine engine installed in a land-based power generating plant is also subjected to high operating temperatures and temperature transients, but it may also be required to operate at full power and at its highest temperatures for very long periods of time, such as for days or even weeks at a time. Prior art insulating systems are susceptible to degradation under such conditions at the elevated temperatures demanded in the most modern combustion turbine systems.
One high temperature coating used to protect jet engine and gas turbine components is Cr.Al.Ti, overlayed with a stabilized zirconia ceramic, as taught in U.S. Pat. No. 5,783,315 (Schaeffer et al.). Many of the ceramic thermal barrier layers are deposited as a columnar structure in the direction of the thickness, as taught in U.S. Pat. No. 4,321,311 (Strangman). This structure can be formed by electron beam physical vapor deposition (xe2x80x9cEBPVDxe2x80x9d) as in Bruce et al. U.S. Pat. No. 5,683,825, or a combination of electron beam deposition and ion beam irradiation, or the like, such as the ZrO2 thermal barrier layer taught in U.S. Pat. No. 5,630,314 (Kojima et al.). Laminates of as many as 20 alternative layers of alumina and yttria stabilized zirconia have also been used as thermal barriers to improve resistance to heat flow for superalloys used in air cooled gas turbines, as taught in U.S. Pat. No. 5,687,679 (Mullin et al.).
Other barrier coating materials include MgTiO3 and Mg2TiO4, as taught in U.S. Pat. No. 5,180,285 (Lau). Padture et al., in J. Am. Ceram. Soc. 80[4]1018-20 (1997) have suggested Y3AlxFe5-xO12, and Y3 Al5O12 (YAG), as well as possibly Y3Al5O12, where Gd, Er or La can substitute for some of the Y sites, as possible thermal barrier coatings.
While zirconia based ceramics provide excellent thermal barrier coatings for a variety of substrates such as turbine blades, more sophisticated coatings are needed to extend the operating temperature capabilities of combustion turbine engines beyond the current state of the art. Recent increases in rotor inlet temperatures are backing up against inherent limitations in the temperature at which yttria stabilized zirconia may be used. Long term exposure of yttria stabilized zirconia above approximately 1200xc2x0 C. can lead to phase destabilization, sintering of the coating, loss of coating compliance and ultimately, possible premature thermal barrier coating failure. Further advances in gas turbine operating temperatures therefore require a ceramic thermal barrier coating capable of surface temperatures in excess of 1000xc2x0 C.
Therefore, it is a main object of this invention to provide improved thermal barrier coating layers for use on underlayers, such as alumina and MCrAlY, protecting turbine components, such as superalloy turbine blade assemblies that can operate over 1000xc2x0 C. for extended periods of time with reduced component degradation.
These and other objects of the invention are accomplished by providing a turbine component containing at least one layer of a thermal barrier coating selected from the group consisting of LaAlO3, NdAlO3, La2Hf2O7, Dy3Al5O12, Ho3Al5O12, ErAlO3, GdAlO3, Yb2Ti2O7, LaYbO3, Gd2Hf2O7, and Y3Al5O12. Usually, the thermal barrier coating will be deposited upon an MCrAlY type alloy layer covering an alloy turbine substrate; where M (xe2x80x9cmetalxe2x80x9d) is selected from the group consisting of Fe, Co, Ni and mixtures thereof. The MCrAlY layer may have an aluminum oxide layer, resulting from oxidation during the deposition coating process or during service. The turbine component can be a turbine blade, a turbine vane, a transition piece, a combustor, and ring segments, or the like, of a high temperature gas turbine, where the thermal barrier coating helps protect the component from impact and erosion by particulates, and from a hostile thermal environment. These coatings can extend the operating temperature capabilities of combustion turbine engines beyond the current state of the art zirconia coatings.